The Revival of interest in flight by man's own muscle power brings back to me some early attempts to solve this problem some 30 years ago.

Needless to say I was—and still am—sold on this idea ever since reading Lilienthal's famous book as a youngster in school. Then came the Wright Brothers and the early years of flying and the years of the First Great War at the end of which I found myself as an Aerodynamicist, designing wing sections and calculating induced drag according to Prandtl's newly discovered wing theory.

The end of the war found us meditating about gliding and the possibility of soaring flight. And so came the Rhoen, and the “Sturm und Drang” of the years on the Wasserkuppe.

Surface pressure measurements were obtained during a three-dimensional vortex interaction with a NACA 0015 aerofoil. The upper and lower surfaces of the blade experienced different aerodynamic loads which appear to be controlled by the impact of the vortex axial core flow on the blade surface. On the upper surface of the blade, where the vortex core flow was away from the aerofoil, the interaction was characterised by the generation of a suction peak. On the lower surface, where the axial component was towards the blade, a pressure pulse developed and seemed to be influenced by the vortex approach angle. These features resulted in rapid changes in normal force and quarter chord pitching moment during the interaction. This impulsive loading of the blade may provide some explanation for sound generation and control degradation problems associated with the tail rotor of helicopters.

To attempt to discuss aerodynamic theory in a single lecture is perhaps a rather rash undertaking, for the subject is so vast that a proper discussion would require a volume, or even a series of volumes. My object, however, is not to cover the whole field of aerodynamic theory with its manifold applications to aeronautical engineering, but rather to discuss some of the fundamental conceptions on which the theory rests, and to discover where theory can be helpful and where it is necessary to rely wholly on experimental results.

The effect of streamwise slots and grooves on a normal shock wave-turbulent boundary-layer interaction has been investigated experimentally at a Mach number of 1.3. The surface pressure distribution for the controlled interaction in the presence of slots featured a distinct plateau. This was due to a change in shock structure from a typical unseparated normal shock wave-boundary-layer interaction to a large bifurcated lambda type shock pattern. Velocity measurements downstream of the slots revealed a strong spanwise variation of boundary-layer properties, whereas the modified shock structure was found to be relatively two-dimensional. Cross flow measurements indicate that slots introduce streamwise vortices into the flow. When applied to an aerofoil, streamwise slots have the potential to reduce wave drag while incurring only small viscous penalties. In the presence of grooves the interaction was initially found to be significantly different. A bifurcated shock structure was observed but the trailing leg appeared stronger and featured a second lambda foot. Oil flow visualisation also revealed differences in the interactions, with the region of suction and blowing being limited to a smaller extent of the grooved control surface. The amount of crossflow present was reduced compared to the slotted control surface. By varying the internal geometry of the grooves it was found that the interaction could be modified to be similar to that in the presence of slots indicating that a more practical control device can be designed.

Wind tunnel data from a tanker wing and receiver aircraft model at varying vertical separation have been compared with theoretical results. In the aerodynamic model the tanker wing is represented by a horseshoe vortex while the aerodynamic loads on the receiver are determined by the vortex lattice method and lifting-line theory, although an approximate method is used to determine the side force on the fin. In the longitudinal case data were obtained for low, mid and high tailplane positions and, with the exception of the pitching moment results, fairly good agreement is obtained between theory and experiment. The relatively small differences are due mainly to the wind tunnel boundary interference effect which could not be quantified for the pitching moment measurements. The lateral aerodynamic interference was determined by banking the tanker wing and displacing it sideways and by yawing the receiver model. Fairly good agreement is obtained between the theory and experiment for the most significant terms which are the rolling moments due to bank and sideways displacement. The effect of the sidewash due to the tanker wake on the receiver in yaw is found to be relatively insignificant. Over the range of bank, yaw and sideways displacements tested the results are almost linear.

The first object of this paper is to show the concordance between three different theoretical approaches to flame stabilisation and to compare published experimental data with the theoretical predictions. Good agreement between the theory and experimental results is obtained. The second object is to show that, at low pressures, the predictions of these theories must be qualified in view of the increasing importance of heat losses from the flame region. It is shown that there must exist a minimum pressure below which a flame cannot be maintained in a given combustible mixture, irrespective of the rate of flow of the mixture, unless heat losses are entirely eliminated.

The future air traffic management (ATM) concept envisaged by the Single European Sky ATM Research – SESAR – and the USA equivalent NextGen, mark a paradigm shift from the current reactive approach of ATM towards holistic strategic collaborative decision making. The core of the future ATM concept relies on common situational awareness over potentially large time-horizons, based upon the user operational intent. This is beyond human capabilities and requires the support of automation tools to predict aircraft state throughout the operation and provide support to optimal decision making long before any potential conflict may arise. This is achieved with trajectory predictors and conflict detectors and resolvers respectively. Numerous tools have been developed, typically geared towards addressing specific airborne applications. However, a comprehensive literature search suggests that none of the tools was designed to predict trajectories throughout the entire operation of an aircraft, i.e. gate-to-gate. Yet, such functionality is relevant in the holistic optimisation of aircraft operations. To address this gap, this paper builds on an existing en route trajectory prediction (TP) model and develops novel techniques to predict aircraft trajectories for the transitions between the ground- and enroute-phases of operation and for the ground-phase, thereby enabling gate-to-gate (or enroute -to-enroute) TP. The model is developed on the basis of Newtonian physics and operational procedures. Real recorded data obtained from a flight data record (FDR) were used to estimate some of the input parameters required by the model. The remaining parameters were taken from the BADA 3.7 model. Performance results using these flight data demonstrate that the proposed TP model has the potential to accurately predict gate-to-gate trajectories and to support future ATM applications such as gate-to-gate synchronisation.

The integral equation approach to the mixed flow problem of infinite wings at high subsonic speeds is adopted for non-circulatory and circulatory (lifting) flows. The solutions are determined from a system of non-linear algebraic equations and, to ensure always unique solutions, the method of differentiation with respect to a parameter has been applied. The resulting Cauchy problem is then solved with the linearised flow solution as the initial value vector. For the case of embedded shocks in the flow field, the method of steepest descent has been added to the calculation scheme. Results for subcritical and supercritical flows past aerofoils are given and compared with solutions obtained by finite-difference techniques.

A numerical investigation has been conducted into the influence of flow compressibility effects around an open-wheeled racing car. A geometry was created to comply with 2012 F1 regulations. Incompressible and compressible CFD simulations were compared – firstly with models which maintained Reynolds number as Mach number increased, and secondly allowing Mach number and Reynolds number to increase together as they would on track. Results demonstrated significant changes to predicted aerodynamic performance even below Mach 0·15. While the full car coefficients differed by a few percent, individual components (particularly the rear wheels and the floor/diffuser area) showed discrepancies of over 10% at higher Mach numbers when compressible and incompressible predictions were compared. Components in close ground proximity were most affected due to the ground effect. The additional computational expense required for the more physically-realistic compressible simulations would therefore be an additional consideration when seeking to obtain maximum accuracy even at low freestream Mach numbers.

The HART II test of 2001 provided a significant and comprehensive database of rotor loads, blade motion, blade pressure distribution, acoustic signature and flow field data. A sub-set of these were released for public access worldwide and an International HART II Workshop was established in 2005. This Workshop, open to the international rotorcraft community, has been held since then at both the AHS Annual Forum as well as the European Rotorcraft Forum. Following the requests of participants, the HART II International Workshop database has been undergoing continuous development. This paper describes the current content and status of the data and the documentation material provided within the workshop database.

This study investigates and proposes a fire detection and suppression system for a smart air cargo container. A series of smoke spread and fire evolution numerical models are executed to assess the performance of container-based fire detection in various fire scenarios. This is to identify the worst case and optimise the location and threshold setting of fire detection sensors, achieving the shortest detection time. It is found that the fire detection threshold (reduction in light transmission = 12%/ft) for a container-based system can be set at three times the standard activation threshold for a cargo-based fire detection system, which can reduce the number of false alarms by three orders of magnitude. Moreover, effectiveness analysis of passive fire protection for the glass fibre-reinforced polymer-made smart container indicates an allowable leakage size of 0.01m2. The risk of internal overpressure has been found to be negligible for the leakage size required by aircraft pressure equalisation.

A project was recently completed that investigated the ability to predict the onset of flutter using tools like the flutterometer. This project used an experiment called the aerostructures test wing that was flown while mounted to the flight test fixture on an F-15. Several flight tests were conducted to expand the envelope and determine the aeroelastic dynamics of the experiment. The final flight ended with destruction of the experiment due to the onset of flutter. The flutterometer attempted to predict this onset by analysing the flight data. The results indicate the flutterometer is able to generate a conservative estimate of the flight conditions associated with flutter. This paper details the flight tests of the experiment and the resulting predictions from the flutterometer.

In a paper in the number of this journal for November, 1925, it was shown that if a flying machine were fitted with wings of standard section and these were flapped in a rhythmical manner, the machine would be supported and propelled, its weight, dimensions and velocity being those of a typical bird; the degree of accuracy attempted in the analysis did not exceed five per cent.

A subsequent examination of the power involved established a high efficiency; but this was due to a large amount of negative work during up–beats; and although such conditions might be maintained in a mechanical model they were unlikely in a bird, which has a large muscle, the pectoralis minor, for lifting its wings. In a bird efficiency seemed to require that the angle of incidence of the outer portion during an up–beat should be negative (see §7, p. 593, of the previous paper); the angle was however made zero (see §9) in order to secure adequate lift, and it is this feature which gives rise to the negative work.

The nature of handling criteria and their uses in relation to the prediction of aircraft handling characteristics are discussed. A brief history is given of the development of ideas on handling, and a survey is made of criteria for modern aircraft, looking particularly at the physical factors affecting the pilot, as an indication of the basis of the criteria. The novel handling features arising with new aircraft species like slender wings and VTOL are considered. The scope of flight simulators and variable stability aircraft for systematic study of handling is discussed and the potentialities of theoretical study by means of servo analysis techniques considered. Finally, possible lines of development in relation to the assessment of aircraft handling characteristics are suggested.

Aerodynamic optimisations of a blended wing-body (BWB) aircraft are presented. A discrete adjoint solver is used to calculate efficiently the gradients, which makes it possible to optimise for a large number of design variables. The optimisations employ either a variable-fidelity method that combines low- and high-fidelity models or a direct sequential quadratic programming (SQP) method. Four Euler optimisations of a BWB aircraft are then presented. The optimisation is allowed to change a series of master sections defining the aircraft geometry as well as the sweep angle on the outer wing for two of the optimisations. Substantial improvements are obtained, not only in the Euler mode but also when the optimised geometries are evaluated using Reynolds-averaged Navier-Stokes solutions. Some interesting features of the optimised wing profiles are discussed.

Experimental results are presented of the measurement of skin friction in fully developed turbulent flow in concentric annuli using Preston tubes situated on the inner and outer annular surfaces. Both Preston's calibration and Patel's calibration were used to evaluate the results. It was found that the latter gave excellent results. Several radius ratios were investigated with a reasonable range of the annulus Reynolds number. The good agreement was not affected by radius ratio or smallness of core tube within the range of these parameters investigated here.

Oil-jet lubrication and cooling of high-speed gears is frequently employed in aeronautical systems, such as novel high-bypass civil aero engines based on the geared turbofan technology. Using such oil-jet system, practitioners aim to achieve high cooling rates on the flanks of the highly thermally loaded gears with minimum oil usage. Thus, for an optimal design, detailed knowledge about the flow processes is desired. These involve the oil exiting the nozzle, the oil impacting on the gear teeth, the oil spreading on the flanks, the subsequent oil fling-off, as well as the effect of the design parameters on the oil flow. Better understanding of these processes will improve the nozzle design phase, e.g. regarding the nozzle positioning and orientation, as well as the nozzle sizing and operation.

Most related studies focus on the impingement depth to characterize the two-phase flow. However, the level of information of this scalar value is rather low for a complete description of the highly dynamic three-dimensional flow. Motivated by the advancements in numerical methods and the computational resources available nowadays, the investigation of the oil-jet gear interaction by means of computational fluid dynamics (CFD) has come into focus lately.

In this work, a numerical setup based on the volume-of-fluid method is presented and employed to investigate the two-phase flow phenomena occurring in the vicinity of the gear teeth. The setup consists of a single oil-jet impinging on a single rotating spur gear. By introducing new metrics for characterizing the flow phenomena, extensive use of the possibilities of modern CFD is made, allowing a detailed transient and spatially resolved flow analysis. Thus, not only the impingement depth, but also the temporal and spatial evolution of wetted areas on the gear flanks, as well as the evolution of the oil volume in contact with the gear flanks are extracted from the simulation data and compared in a CFD study.

The study consists of 21 different simulation cases, whereby the effect of varying the jet velocity, the jet inclination angle, the jet diameter, and the gear speed are examined. Consistent results compared to a simplified analytical approach for the impinging depth are obtained and the results for the newly introduced metrics are presented.

An analytical solution, using complex variable methods, is given for the problem of the stress distribution due to an elliptical hole, reinforced around its boundary, in a plane sheet subjected at infinity either to an arbitrary constant stress system or to a bending type stress system. Numerical results were obtained for a wide range of parameters, including three different shapes of ellipse, and ten different amounts of reinforcement. Poisson's ratio was assumed to be 1/3.

An integrated artificial immune system-based scheme that can operate over extended areas of the flight envelope is proposed in this paper for the detection and identification of a variety of aircraft sensor, actuator, propulsion, and structural failures/damages. A hierarchical multi-self strategy has been developed in which different self configurations are selected for detection and identification of specific abnormal conditions. Data collected using a motion-based flight simulator were used to define the self for a wide area of the flight envelope and to test and validate the scheme. The aircraft model represents a supersonic fighter, including model-following direct adaptive control laws based on non-linear dynamic inversion and artificial neural network augmentation. The proposed detection scheme achieves low false alarm rates and high detection and identification rates for all the categories of failures considered.

A calculation procedure for modelling the interaction between shock waves and attached or separated turbulent boundary layers is introduced. The numerical framework, applicable to general curved grids, combines cell-vertex storage, a Lax-Wendroff time-marching scheme and multigrid convergence acceleration. The main numerical ingredients of the procedure are documented in some detail, with particular emphasis placed on the inclusion of viscous and turbulence transport within the cell-vertex framework, which was originally formulated for inviscid flow. An investigation of the predictive performance of alternative transport models of turbulence has been the primary objective of the present work. Particular attention has been focused on a comparison between variants of low Reynolds number k-ε models and an algebraic variant of a Reynolds-stress transport closure in strong interaction situations, including shock-induced separation. The turbulence models are introduced, and important numerical issues affecting their stable implementation are discussed. The calculation procedure is then applied to two confined transonic flows over bumps — one incipiently and the other extensively separated (Delery Cases A and C) — and to the transonic flow around the RAE 2822 aerofoil at two angles of incidence (Cases 9 and 10). The investigation demonstrates that the eddy-viscosity models tend to seriously underestimate the strength of interaction, particularly when separation is extensive. The performance of the Reynolds-stress model is not entirely consistent across the range of conditions examined. In the case of bump flows, the model displays strong sensitivity to the shock, predicting excessive boundary layer displacement in Case A, a broadly correct separation pattern in Case C and insufficient rate of post shock recovery in both cases. The aerofoil flows are either attached or incipiently separated, and the benefits arising from Reynolds-stress modelling are modest. Neither the k-ε model nor the Reynolds-stress closure is able to return a satisfactory representation of the most challenging RAE 2822 Case 10; at least not with the recommended windtunnel corrections for freestream Mach number and angle of incidence.